(1) Field of the Invention
The present invention relates to a cooling arrangement for use in a turbine engine component.
(2) Prior Art
FIG. 1 illustrates a current cooling scheme for a turbine blade 10. It consists of a hybrid application of embedded microcircuit panels 12 running axially along the airfoil walls 14 and 16 in combination with a set of film cooling holes. The airfoil active convective cooling is done through a series of microcircuits 12 in the mid-body and trailing edge portions of the airfoil 18, supplemented with film cooling by a series of film holes 20. There are two considerations with this blade that could be improved upon. First, the axial circuits do not take full advantage of pumping; therefore, dedicated feed cavities are used for independently feeding each circuit. This leads to an increased number of airfoil ribs 22. Second, as a result, the ribs 22 are relatively cold when compared with the outer layers of the airfoil walls.
As the blade 10 ramps up in load, the airfoil outer layers experience relatively hot metal temperatures. If the temperature is sufficiently high, a stress relaxation process occurs at these airfoil locations, leading to relatively high strains (deformations). Simultaneously, the relative cold inside ribs 22 experience an increase in stress as the load to the part needs to be shared by the entire airfoil 18. This balance in the stress-state of the airfoil occurs every time a blade is ramped up, causing some amount of irreversible damage, which, in excessive limits, can lead to catastrophic failures. If these limits are not approached, the amount of damage accumulation can take some time or cycles. That is, long enough to make the design viable for the require life targets. Two modes of failure exists: (a) creep; and (b) fatigue. Oxidation also occurs, but is not discussed as it can be incorporated in creep damage due to the reduced load-bearing capability from metal-oxide attack. The creep damage is related to blade temperature; but fatigue is related to temperature differences in the blade, in particular, the outer relative hot airfoil layers and cold internal ribs. It is therefore desirable to reduce the outer metal temperatures, and the thermal gradients in the part.